1. Field of the Invention
This invention relates broadly to rockets. More particularly, this invention relates to ignition systems for hybrid and solid rocket motors.
2. State of the Art
Rocket motors generally fall into three classes: solid propellant motors in which a solid fuel element undergoes combustion to produce thrust that propels the rocket, liquid propellant motors that accomplish the same function with a liquid fuel material, and hybrid rocket motors. Hybrid rocket motors may be characterized as a cross between a solid propellant motor and a liquid propellant motor. A hybrid motor generally uses a propellant consisting of a fluid oxidizer and a solid fuel element; however, they may use a combustible liquid fuel and a solid oxidizer.
Both solid rocket motors and hybrid rocket motors use an ignition system to initiate propellant combustion by creating a flame source in the combustion chamber of the rocket. The combustion chamber in a solid rocket motor houses the propellant, whereas in a hybrid rocket motor the combustion chamber typically houses solid fuel and the fluid oxidizer is fed into the chamber from a tank. Combustion of the solid or hybrid propellant generates thrust as the high pressure combustion products are discharged through the rocket nozzle.
Referring now to Prior Art FIG. 1, a prior art hybrid rocket 910 is shown. The rocket 910 generally includes a combustion chamber 912 provided with a solid fuel grain 914, a main oxidizer tank 916 adapted to feed an oxidizer 917, e.g., nitrous oxide, into the combustion chamber 912 through a valve 918, an aft nozzle 920, and a forward nose cone 922.
An ignition system 923 is provided for initiating combustion of the propellant. The ignition system 923 includes a rigid metal tube 924 axially extending into the combustion chamber 912. The tube 924 includes a longitudinal opening 926 and optionally a single set of one or more radial openings 928. The ignition system 923 also includes a outboard tank 930 dedicated to the ignition system, a tank valve 932, a regulator 934 and a low pressure solenoid valve 936. The tank 930 is provided with a pressurized fluid oxidizer 931, such as gaseous oxygen at 3000 psi. The tank valve 932 controls release of the oxidizer 931 from the tank 930. The regulator 934 controls the pressure of the oxidizer 931 after the oxidizer is released from the tank 930 and preferably drops the pressure down to approximately 100 psi. This lower pressure prevents the oxidizer 931 from “blowing out” the ignition flame, discussed below. The solenoid valve 936 controls release of the oxidizer 931 into the tube 924 and up into the combustion chamber 912. The ignition system 923 further includes two wires 938, 940 having exposed leads 942, 944 situated outside the tube 924, adjacent the openings 926, 928 of the tube 924 and near the forward bulkhead 946 of the rocket motor, as well as an ignition source, such as a neon sign transformer 948 capable or producing 10,000 V at 30 mA.
In operation, the tank valve 932 is opened and the solenoid valve 936 is actuated to allow the lower pressure oxidizer 931 to fill the chamber 912. Substantially simultaneously (e.g., within a few milliseconds), the transformer 948 is activated to create a high voltage arc across the leads 942, 944, which operates as the spark for ignition. The solid fuel grain 914 becomes the fuel source for ignition, as the lower pressure oxidizer 931 reacts with the exposed surface of the fuel grain 914. The lower pressure oxidizer is continually fed into the motor, preferably until the entire surface of the hybrid fuel grain is lit and the decision is made to open the main oxidizer tank valve 918 for the main propulsion phase. The arc causes the ignition oxidizer 931 to combust with the solid fuel grain 914 and thereby creates a flame. Such an ignition system is described in more detail in U.S. Pat. No. 5,715,675 which is hereby incorporated by reference herein in its entirety.
This type of ignition system has several shortcomings. First, in general, the hybrid fuel grain port is significantly larger in diameter, i.e., across 950, than the largest low pressure oxidizer tube that can be advanced up through the throat 952 of the nozzle 920. As a result, the ignition source and oxidizer are not in direct contact with the fuel grain. This creates an ignition delay while the spark “jumps the gap” toward the fuel grain. Second, the metal tube upon lift-off is blown out the nozzle. This forceful contact of the metal tube against the nozzle can result in damage to the nozzle and prevent successful operation of the rocket.